Aircraft supplemental thrust device and method of operating the same

ABSTRACT

A supplemental thrust system for an airplane comprising a submerged inlet duct, a diffuser section in communication with the inlet duct, and a duct outlet in communication with the diffuser section, an outlet nozzle in communication with the duct outlet, a turbocharger unit comprising one or more turbocharges in communication with each other to provide compressed air to the engine and the cabin, a heat exchanger unit comprising a plurality of heat exchangers in the diffuser section to provide cooling liquid to cool pressurized air in the cabin, engine intake air and the engine jacket, where each turbocharger is coupled to the intake manifold of an engine to increase the power of the engine by introducing compressed air into the manifold.

CROSS REFERENCE TO RELATED APPLICATIONS

This application claims priority to U.S. Provisional Patent ApplicationSer. No. 61/753215, filed Jan. 16, 2013, which is incorporated byreference herein in its entirety.

BACKGROUND OF THE INVENTION

The present invention was borne out of frustration with the cost andinefficiency of the airlines' hub-and-spoke transportation model. Thismodel was conceived by the airline industry, initially in an attempt torestrain passengers from using interline transfers to arrive at theirdestinations. It requires dense concentrations of passengers both at therelatively few hub facilities and in ever larger aircraft flying tofewer and fewer destinations. The inefficiencies for the traveler ariseout of the time wasted traveling long distances from their true originto the large hub or major airport, enduring the lengthy lines atcheck-in and security check points, and the ever-longer boarding processon the ever larger aircraft. In addition, the traveler must often fly tocities that are well out of the way to his final destination, andtransfer with additional wasted connection times. The result is that forshort trips (approximately 500 miles) average speeds reduce to thevicinity of 100 mph, and many longer trips that involve just oneconnection drop to 200 to 300 mph average. This inefficiency raisescosts for the consumer, especially where the inefficiencies requireovernight stays in order to catch connecting flights. There is anadditional factor which is a disadvantage of the current hub and spokesystem. The current system creates large concentrations of people, bothat terminals and in ever larger aircraft, that create prime targets forterrorist activity. Larger numbers of much smaller aircraft operating ina widely distributed transportation system would present a moredifficult target for any significant military or terrorist activity.

Clearly, there are compelling reasons for wanting an air transportationsystem that is economically superior to our current air transportationsystem in acquisition, operation and maintenance costs. To be a viablecompetitor, the system should have true origin to true destinationspeeds that significantly exceed current system speeds. It shouldrequire no additional infrastructure, and it should package passengersin small enough units that both the passenger load and the aircraft aremilitarily insignificant targets. To be truly competitive, it shouldprovide non-stop transcontinental and intercontinental travel from anylocal airport to any other local airport. And ticket prices should behighly competitive with current average ticket prices.

Such a transportation system requires a unique aircraft. It must becapable of operation from any current airfield. Preferably, it wouldhave operating costs well below current costs and competitive withcommercial airliners, cruise at higher system speed than currentcommercial aircraft, have a longer range with full passenger and luggageload than most current business aircraft, provide passenger comfortcomparable to commercial aircraft, and be capable of all-weatheroperation. The plane should also provide for ease of maintenance andrequire only a single pilot.

SUMMARY OF THE INVENTION

One embodiment consistent with the present invention includes asupplemental thrust system for an airplane comprising a submerged inletduct, a diffuser section in communication with the inlet duct, and aduct outlet in communication with the diffuser section, an outlet nozzlein communication with the duct outlet, a turbocharger unit comprisingone or more turbocharges in communication with each other to providecompressed air to the engine and the cabin, a heat exchanger unitcomprising a plurality of heat exchangers in the diffuser section toprovide cooling liquid to cool pressurized air in the cabin, engineintake air and the engine jacket where each turbocharger is coupled tothe intake manifold of an engine to increase the power of the engine byintroducing compressed air into the manifold.

In another embodiment, the turbine exhaust of each of the turbochargerson each side converge into a single duct that is connected to the ductoutlet.

In another embodiment, an expansion duct is positioned between the ductoutlet and the outlet nozzle.

In another embodiment, the inlet duct is a NACA duct on the exterior ofan aircraft.

DESCRIPTION OF THE DRAWINGS

Details of the present invention, including non-limiting benefits andadvantages, will become more readily apparent to those of ordinary skillin the relevant art after reviewing the following detailed descriptionand accompanying drawings, wherein:

FIG. 1A depicts one embodiment of an aircraft consistent with thepresent invention;

FIG. 1B depicts a breakaway view of the aircraft of FIG. 1;

FIG. 1C depicts a rear perspective view of the rear fuselage of FIG. 1A;

FIG. 2 shows a top perspective view of the truss element;

FIG. 3 depicts a breakaway view of the aircraft including the pressurevessel;

FIG. 4 depicts one embodiment of one of the plurality of standoffs usedto secure the pressure vessel;

FIG. 5 depicts the attachment of skin to the truss elements;

FIG. 6A depicts the front landing gear affixed to the front bulkhead;

FIGS. 6B-6E depict the front landing gear retracting into the frontfuselage;

FIG. 7A depicts the main landing gear connected to truss element;

FIGS. 7B-7E depict the main landing gear retracting into the rearfuselage;

FIG. 8 depicts a heat recovery system used to increase the efficiency ofthe aircraft;

FIG. 9 depicts a side view of the wing spar of the aircraft of FIG. 1;

FIG. 10A depicts a flap control system included in the wing of theaircraft in FIG. 1;

FIG. 10B depicts the flap control system with the plates removed;

FIG. 11A depicts the flap control system extending to lower the flaps;

FIG. 11B depicts the flap control system extending the flap downward;

FIG. 11C depicts the flap control system as it extends further outwards;

FIG. 11D depicts the flap control system with the foreflap and flap inthe full extended position;

FIG. 12 depicts the spoiler actuation system used to actuate the spoilerof FIG. 9;

FIG. 13A depicts a trim actuator that is mechanically coupled to theelevator control system and similarly used in the dorsal fin controlsystem; and

FIG. 13B depicts an interior view of the actuator along the lines A-A.

DETAILED DESCRIPTION OF THE INVENTION

The purpose and advantages of the present invention will be set forth inand apparent from the description that follows, as well as will belearned by practice of the invention. Additional advantages of theinvention will be realized and attained by the methods and systemsparticularly pointed out in the written description and claims hereof,as well as from the appended drawings. The term “top portion” is usedherein to mean the portion of the fuselage farthest from the ground whenthe airplane is not in flight and the term “bottom portion” is herein tomean the portion of the fuselage closest to the ground when the airplaneis not in flight.

The design of the present invention makes use of aerodynamic shapes thatare extensively laminar within their Reynolds number operating regime.Intersections of wing, empennage and fuselage are minimized, ellipticallift profiles are used on all lifting surfaces, and wing and horizontaltail shapes are approximately elliptical. The fuselage shape is derivedfrom a modified zero camber extensively laminar airfoil section revolvedabout the longitudinal axis, thus making full use of pressure recoveryto minimize form drag. The external aerodynamic shapes are mostlyprovided by gloves that fit over the frame of the aircraft, but areisolated from the frame so as to reduce surface waviness under load toan absolute minimum. This also permits easy one piece complete removalof the external skins for inspection of the frame and frame elements andmaintenance of the operating systems attached to the frame.

The wing structure of the aircraft consists of a box-and-channelstructure that extends across approximately 90% of the span of the wingstructure and is open to the rear but stabilized in compression. Thewing structure is a composite beam with ply orientation and shapetailored to provide structural coupling in bending and torsion togenerate variable wing washout as a function of bending to limitvertical wing loading and to provide damping of the major flutter modes.Tail surfaces have similar spar-and-glove design to allow for ease ofinspection of all primary structure, decoupling of structuraldeformation from skin surfaces, and ease of exchange of external skinwith new shapes for rapid repair of damaged surfaces as well as exchangeof airfoil shapes with updated shapes or different internal systemsshould they become available.

FIG. 1A depicts one embodiment of an aircraft consistent with thepresent invention. The aircraft includes a forward fuselage 1, a rearfuselage 2, a midwing 3, a vertical fin 4, a ventral fin 5, a horizontaltail 6 and a pusher type propeller 7. The forward fuselage 1 and rearfuselage are covered in an external skin. The external skin may be madeof a rigid fiber reinforced composite or metal such as, but not limitedto, an aluminum alloy such as aluminum 2024 or aluminum 7078 or anyother rigid material meeting a maximum waviness tolerance of 0.001inches per inch measured over a two-inch span.

FIG. 1B depicts a breakaway view of the aircraft of FIG. 1. The frame ofthe aircraft includes a forward bulkhead 8 connected to an upper truss 9on one end and two lower forward trusses 10 and 11 on an opposite sideof the forward bulkhead 8. The truss elements 9, 10 and 11 may be boxtype truss structures where the ends of the truss elements 9, 10 and 11taper towards the forward bulkhead 8, providing improved stiffness atthe intersection of the truss elements 9, 10 and 11 and the forwardbulkhead 8. The truss elements 9, 10 and 11 are made of a rigid materialincluding metal, fiberglass including S glass, or an equivalentmaterial. Each composite truss element 9, 10 and 11 also includes aunidirectional upper cap, a unidirectional lower cap andshear/compression panels connecting the upper and lower caps. The shearpanels may be comprised of +45/−45/0/90 plies of fiberglass, such as Sglass or equivalent, configured for crush stiffness when loaded invertical compression and for the minimal shear loading required by thetriangulated configuration of the upper and lower caps.

Each truss elements 9, 10 and 11 extends from the forward bulkhead 8 tothe main bulkhead 15 where the truss elements 9, 10 and 11 are affixedto the main bulkhead 15 by fastener devices 12, 13 and 14. The fastenerdevices 12, 13 and 14 may be comprised of transverse beams which may beformed of metal or a composite such as carbon fiber. Each fastenerdevice 12, 13 and 14 is affixed to a respective truss element 9, 10 and11 by a securing device such as a bolt passing through the fasteningdevice 12, 13 or 14, the corresponding truss element 9, 10 and 11 and aportion of the main bulkhead 15. Each fastening device 12, 13 and 14 isattached to its respective truss element 9, 10 or 11 by wrapping theinner and outer plies of fastening device 12, 13 or 14 around the trusselements 9, 10 or 11 and doubling those plies back upon their outer andinner mating plies, respectively, thus mechanically locking thefastening device 12, 13 or 14 to respective truss element 9, 10 or 11.Similar mechanical locking is used on the truss elements 19 and 20 ofthe rear fuselage. A main bulkhead transverse beam 16 is affixed to theexposed portions of the periphery of the main bulkhead 15 and isconnected to the truss elements 9, 10 and 11.

Truss element 19 is affixed to the top portion of the main bulkhead 15such that the central axis of the truss element 19 is substantiallyco-linear with the central axis of the truss element 9. Truss element 20is affixed to the bottom of the main bulkhead 15, and truss elements 21and 22 are affixed to opposing sides of the main bulkhead 15. Each ofthe truss elements 19, 20, 21 and 22 may be box type beams. Trusselements 21 and 22 are configured to resist lateral loads induced by thevertical fin 5 and to provide support for skin cutouts required for themain landing gear doors and upper access hatches as described in furtherdetail herein.

Truss elements 19 and 20 extend from the main bulkhead 15 to a rear tailcone 24. Each truss element 19 and 20 is affixed to the rear tail cone24 using any known method of connection such as bolts, rivets orbonding. The upper surfaces, the surfaces facing away from the centerportion of the aircraft, are coplanar with the surface of the tail cone24. The truss elements 21 and 22 are each affixed to a rear traversebulkhead 25, shown in FIG. 1C, and to a forward traverse bulkhead 26. Abox section support 27, shown in FIG. 1C, is positioned between the reartraverse bulkhead 25 and forward traverse bulkhead 26 on the tail cone24 to provide support for a vertical fin spar 28. A horizontal tail spar29 is positioned between the rear bulkhead 25 and an elevator bulkhead34, shown in FIG. 1C.

A fuel tank 33 is positioned adjacent the main bulkhead 15 in the rearfuselage 2. The fuel tank 33 may be semicircular in shape and bepositioned above the mid wing 3. The fuel tank 33 is a separatereplaceable bladder manufactured of a metal lined, highly damagetolerant composite structure that is internal to the fuselage andmounted on top of the wing spar, and is outside of the pressure vessel.Conventional wing tanks are difficult to seal and drain, and they arehighly vulnerable to rupture in a crash due to their exposed distributedlocation along the wing span. With wing tanks, volumetric rearrangementin the event of crash-induced high G force loading is difficult toaccomplish due to the walls of the tankage being part of the primarystructure of the wing. By separably mounting the tank above the heaviestprimary structure in the center of the aircraft, and by using amoderately volume-inefficient shape, volume rearrangement and thussurvivability of the tank is enhanced.

FIG. 1C depicts a rear perspective view of the rear fuselage 2 of FIG.1A. The mid wing 3 is coupled to the main bulk head 15 by the sleeve 17.The sleeve 17 is affixed to the main bulkhead 15 by a plurality ofstraps 39. The straps 39 may be made of unidirectional fiberglass suchas S glass, or any other material capable of securing the sleeve 17 tothe bulkhead 15. Each strap 39 extends around the periphery of thesleeve 17 such that a first portion of the strap 39 is in direct contactwith the top surface of the sleeve 17, a second portion of the strap 39is in direct contact with a side surface of the sleeve 17, and a thirdportion of the strap 39 is in direct contact with a lower portion of thesleeve 17. A first end and second end of each strap 39 is affixed to themain bulkhead 15 by any known method of attaching a strap to a bulkhead,including rivets, bolts or bonding.

A gusset 40 is attached to the lower portion of the sleeve 17 on one endand the main bulkhead 15 on the opposite end. The gusset 40 may betriangular in shape, with the wider portion of the gusset 40 connectingto the main bulkhead 15 and the narrower portion of the gusset 40connecting to the bottom surface of the sleeve 17. The gusset 40 acts totransfer upward loading force of the fuselage to the main bulkhead 15.After installation, the fuel tank 33 shown in FIG. 1B may be positionedon the top surface of the sleeve or on a separate horizontal panel oftransverse beam 37, bonded to the structure.

A transverse beam 37 is positioned on the bottom side of each trusselement 21 and 22 and the side surface of the sleeve 17. Half supportring 18 extends from the top surface of the transverse beam 37 adjacentto the truss element 21 to the top surface of the transverse beam 37adjacent the lateral element 22. The top surface of the half supportring 18 is substantially coplanar to the top surface of the trusselements 19, 20 and 21. Full support ring 38 extends from one side ofthe truss element 20 to the opposite side of the truss element 20 suchthat the full support element connects to the truss elements 19, 21 and22. The top surface of the full support ring 38 is substantiallycoplanar with the top surfaces of the truss elements 19, 20 and 22. Eachsupport ring 18 and 38 is attached to truss elements 19, 20, 21 and 22by multi-ply tabs as previously discussed or by any other method ofattaching a support ring to a truss. Additional full and half supportrings may be provided and affixed to the structure in a manner similarto the attachment of the half support ring 18 and full support ring 38.

The horizontal tail spar 29 is affixed between the rear bulkhead 25 andthe elevator bulkhead 34. The horizontal tail spar 29 is a continuoussingle piece spar that is pivotally attached to the rear fuselage by apair of bearing units 36 mounted in a bearing carrier 35. The outersides of the bearing carrier 35 are affixed to the rear bulkhead 25 andthe elevator bulkhead 34. A tail wheel gusset 30 may be connected to thebottom surfaces of the rear bulkhead 25 and elevator bulkhead 34 toprovide ventral fin and propeller protection from a tail strike due toover rotation during takeoff or landing. A wheel extension arm 31 andwheel 32 are rotatively affixed to one end of the gusset 30. An actuatorunit 33 is affixed to the bottom surface of the nose cone 24 between thegusset 30 and the end of the cone 24 such that the wheel extension arm31 and wheel 32 can be extended during and retracted during flight.

FIG. 2 shows a top perspective view of the truss element 20. Trusselement 20 includes a forward portion 552, a rear portion 554, supportunits 404 and a bulkhead connection plate 550. The forward portion 552and rear portion 554 are joined at center joint 560 and the supportunits 553 are affixed to the sides of the truss element 20 at the centerjoint 560. The forward portion 552 and rear portion 554 are connectedsuch that the top surface of the forward portion 552 and the top surfaceof the rear portion 554 form angle theta. In one embodiment, theta isapproximately 180 degrees. In another embodiment, theta is betweenapproximately 150 and approximately 178 degrees.

The truss element 20 has a box structure with four sides and a hollowcenter portion. Openings may be cut along the sides of the truss element20 to reduce the overall weight of the truss element 20 while alsoproviding support for lateral and vertical loads encountered in flight,landing and takeoff conditions. The support units 553 extend from thesides of the truss element 20 at an angle beta relative to the topsurface of the truss element 20. Each support unit 553 includes aconnection plate 410 on the end of the support unit 404 furthest fromthe truss element 20. The bulkhead connection plate 550 is affixed tothe front surface of the truss 20. The bulkhead connection plate 550includes a substantially arc shaped portion that is shaped to engage alower portion of the main bulkhead 15 using connection openings 551. Aplurality of sidewall connection openings 555 are positioned along thesidewalls of the truss element 20 for connecting a motor mount to thetruss element 20.

FIG. 3 depicts a breakaway view of the aircraft including the pressurevessel. The pressure vessel 43 is positioned in the forward fuselageassembly 42 between the main bulkhead 15 and the nose of the aircraft.Because of the differing forms of the loads induced by local loading bypayloads, aerodynamic loads and ground loads and the distributed loadingfrom pressurization, payload-induced loading is applied to the fuselagetruss elements 9, 10, 11, 19, 20, 21 and 22, and not the pressure vessel43, which is isolated from the truss elements 10, 11, 19, 20, 21 and 22.Isolating the pressure vessel 43 eliminates waviness of the externalskin due to pressure deflections as would be the case with aconventional monocoque aircraft fuselage structure. Minimal waviness isa necessary criterion for the maintenance of laminar flow over thefuselage, resulting in corresponding low parasite drag of the fuselage.

The pressure vessel 43 is positioned in the forward fuselage assembly 42such that it is surrounded by the truss elements 9, 10, and 11 and mainbulkhead 15. The pressure vessel 43 is structurally isolated from thetruss by padding rings on the truss elements 9, 10 and 11 that supportthe pressure vessel 43. Vertical deflection of the truss elements 9, 10and 11 will not couple to the pressure vessel 43, and as a consequencestructural loading of the elements 9, 10 and 11 by payloads will produceessentially no induced loads in the pressure vessel 43. Similarly,pressurization of the pressure vessel 43 will contribute no loading tothe truss elements 9, 10 and 11 in any direction because the twostructures are completely decoupled via the pads. The pressure vessel 43is indexed to the truss elements 9, 10 and 11 by a single standoff (notshown) that penetrates the pressure vessel 43 through a close tolerancehole and is sealed to internal pressure of the pressure vessel 43 by acircular seal that is free to slide in the radial direction on thestandoff. The indexing standoff (not show) is one of a number ofstandoffs that penetrate the pressure vessel 43 through oversizedreinforced holes in the pressure vessel 43 and which carry the loadssustained by the floorboards, internal panels and other internalappurtenances through the pressure vessel 43 outwards into the trusselements 9, 10 and 11. All but two of these reinforced holes are looselongitudinal and circumferential fits to the standoffs to allow forpressure vessel expansion, and thus there is only a single longitudinaland circumferential locating position.

The parts of the pressure vessel 43 forward and aft of an index positionare free to expand and contract longitudinally, circumferentially andradially without coupling any loads or deflections into the trusselements 9, 10 and 11 and conversely, truss element deflections cannotproduce induced loading in the pressure vessel 43. The front dome of thepressure vessel 43 is an ideal hemispherical shape with cutouts for awindshield and windows. Those cutouts are ring and strap reinforced toresist the tangential pressure loads, and the panes are coupled to thevessel 43 in only a radial direction. Therefore, no circumferentialloads are transmitted.

The differential thermal expansion and the pressure-induced diaphragmdeflections of the panes from the pressure vessel 43 are also reduced bythe ring and strap reinforcement. In contrast, the doors are setcoplanar to the pressure vessel 43 walls and are fastened in atangentially load bearing semi-continuous fashion to the walls of thepressure vessel 43 around their entire circumference by means of thesealing device 67. Internal pressure increases latching forces of thedoors to the walls of the pressure vessel 43. The doors are thusload-bearing elements of the pressure vessel 43.

FIG. 4 depicts one embodiment of one of the plurality of standoffs usedto secure the pressure vessel. The standoff includes two loaddistribution plates 63 and 69. The external plate 69 is affixed to atruss element 9, 10 or 11. The interior plate 63 is affixed to a loadbearing structure within the pressure vessel 43. A cylindrical standoff64 has opposing ends fastened to the distribution plates 63 and 69 byfasteners 70. The fasteners 70 are configured to carry the full loadapplied to the standoff, and are held in positioned by a lockingmechanism such as a tabbed washers, safety wire or any other means oflocking the fasteners 70 in place. The cylindrical standoff 64 extendsthrough an opening in the wall of the pressure vessel 62. The opening inthe wall of the pressure vessel 43 is sized to accommodate the expansionand contraction of the pressure vessel 43, and the movement of thepressure vessel 43 during operation of the aircraft. Two standoffs 64that are diametrically opposed, are connected to openings in thepressure vessel 43 that do not compensate for expansion and contractionof the pressure vessel 43 during operation.

The openings in the pressure vessel 43 are reinforced by a plate 65 thathas a surface coplanar to the outer surface of the pressure vessel 43.The plate 65 may be made of any material capable of withstandingtangential loads of the pressure vessel 43 including steel, aluminum andalloys thereof, carbon fiber or any other material that can withstandthe tangential loads of the pressure vessel 43. The material of theplate 65 also has thermal expansion and elastic characteristicscomparable to the material used in the pressure vessel 43. In oneembodiment, the pressure vessel 43 and the plate 65 are made from thesame material. The interior portion of the plate 65 engages a washer 66.The washer 66 includes a cylindrical boss sized to accommodate a sealingdevice 67, such as an O-Ring. The sealing device 67 engages thecylindrical standoff 64 such that the washer 66 is in direct contactwith the cylindrical standoff 64. A spring 68 positioned between theplate 69 and the washer 66 forces the washer 66 against the plate 65.

The cylindrical standoffs 64 penetrate the pressure vessel 43 throughthe openings in the pressure vessel 43 wall which reinforced by thewasher 66-spring 68 combination to carry the tangential pressure inducedloads. The standoffs 64 are fastened to truss elements 9, 10 and 11 asnecessary for load distribution. The standoffs 64 are pressure sealed tothe wall of the pressure vessel 43 by means of the washers 66 and spring67, which bosses are sealed by the sealing device 67 that seals thewashers 66 to the cylindrical standoffs 64 by the washers' 66 flat butflexible surface resting on the corresponding flat surfaces provided onthe inside of the wall of the pressure vessel 43. The combination washer66 and spring 67 are free to slide both on the standoff 64 outerdiameter and on the flat on the inside of the pressure vessel 43 wall.The internal diameter of each opening is large enough with respect tothe outer diameter of the penetrating standoffs 64 to allow for allanticipated expansion and contraction of the pressure vessel 43 anddeflections of the truss under load. Using these techniques, thepressure vessel 43 sees only well distributed loading due to internalpressure and is completely isolated from payload-induced loads and otherflight and ground loads. The weight of the pressure vessel 43 itself issupported by elastomeric foam attached to the interior surfaces of thebeams of the forward truss elements 9, 10 and 11. This provides only apadded resting surface for the exterior of the wall of the pressurevessel 43. The pressure vessel 43 can be installed and removed from theforward fuselage 41 as a unit. This is done by separating the forward 41and rear 42 halves of the fuselage and inserting or removing thepressure vessel through the rear opening of the forward fuselage.

The internal dimensions of the forward fuselage truss elements 9, 10 and11 are slightly larger than the maximum pressurized diameter of thepressure vessel 43. The truss elements 9, 10 and 11 are bonded to theexterior skin of the aircraft, and the skin forms a shear web betweenthe top truss element 9 and the bottom truss elements 10 and 11. Thetruss elements 9, 10, and 11 are bonded to the forward bulkhead 8 in atriangulated fashion, and the forward bulkhead carries the nose gearloads into the truss elements 9, 10 and 11. By using multiple standoffpenetrators to carry the loads from inside the pressure vessel 43, tothe truss elements 9, 10 and 11, a relatively uniformly distributed loadon the truss elements 9, 10 and 11 is achieved. This minimizes localdeflections and high stress points that could induce undesirablewaviness into the outer skin of the fuselage. Both the floorboardstructure and the box beams that form the bottom elements of the trussare used as crush structure to manage energy absorption to enhancecrashworthiness. The overall aircraft structure is designed for 26 gultimate loading.

The external skin of the forward fuselage is composed of a formedsandwich panel which is bonded to the truss elements 9, 10, and 11, theforward bulkhead 8 and an attachment ring at the rear of the forwardfuselage. The rear fuselage skin is similar and is bonded to the upper,lower, and side truss elements 19 and 20. The rear half of the fuselagecontains the main bulkhead 15, which is bonded to the forward ends ofthe truss elements 19, 20, 21 and 22 and the rear skin. The sleeve 17 isbonded to the main bulkhead 15 and to two truss elements 21 and 22 whichare likewise bonded to the skin and to the main bulkhead 15. The trusselements 21 and 22 are provided to stiffen the rear fuselage in thelateral direction. This is necessary due to the large skin cutouts forthe main landing gear doors and other access hatches.

The truss elements 19 and 20 are single box beams on both top andbottom. All four box beams and the rear fuselage 43 skin are bonded tothe tail cone 24 which carries the horizontal and vertical tail surfaceattachments and bearings. To allow for a sliding seal surface betweenthe two halves of the horizontal tail and the fuselage, the tail cone 24is surrounded by a removable, mechanically-fastened fairing that iscontoured to fit the rotational movement of the inner surfaces of thehorizontal tail. This fairing is a replaceable wear surface thatprovides the sealing surface for the sliding seal between the horizontaltail and the fuselage.

FIG. 5 depicts the attachment of skin to the truss elements. The forwardskin 44 is bonded to a box ring 49 with a core 50. The rear skin 45 isbonded to the main bulkhead 15, the main bulkhead 15 includes a forwardskin 47, a rear skin 48, and a core 46. Doubler plies or metal doublers51 and 52 provide stress distribution of the local loading generated bythe fasteners, 55. There are a multiplicity of fasteners distributedcircumferentially around the box ring 49 to provide a semi-continuousengagement between the forward skin 44 and the rear skin 45. Thefasteners 55 are shoulder bolts that provide shear coupling between theskins, as well as adequate tensile coupling. The fasteners 55 arethreaded into a sealed nut plate 53 with a counter bored section toengage the shoulder of the fastener 55. To prevent crushing of the coreof the main bulkhead 15, a tubular standoff is bonded to the forwardskin of the bulkhead, 47 and the rear skin of the bulkhead 48. Thisallows the fastener 55 to load the forward bulkhead skin 47 against therear doubler 52 the rear fuselage skin plies 45 the box ring plies 49the forward fuselage skin plies 44 and the forward doubler 51 stacked inthat order without crushing the main bulkhead core 46 or the box ringcore 50.

FIG. 6A depicts the front landing gear 43 affixed to the front bulkhead8. The landing gear 43 may be an oleo type trailing link landing gear.FIG. 6B-6E depict the front landing gear retracting into the frontfuselage. FIG. 6B shows the landing gear 43 in the fully extendedposition. The front landing gear 43 includes an actuation device 612, awheel 602, a swing arm 604, a forward link arm 606, a horizontal hingedplate 608 and an oleopneumatic cylinder 610. The swing arm 604 includestwo parallel plates with one end of each plate being connected to thewheel 602 by an axle that passes through the center of the wheel 602 andthrough corresponding openings in the plates of the swing arm 604. Theother end of the swing arm 604 opposite the wheel 602 is rotativelycoupled to the forward link arm 606 by a pin 610 that allows the swingarm 604 to rotate relative to the forward link arm 606.

The hinged plate 608 is rotatively coupled to the bulkhead 8 by hinges612 connected to the bulkhead 8 such that the plate 608 is pulledtowards the bulkhead 8 as the landing gear 43 is moved to the refractedposition and the plate 608 is moved to a position substantiallyperpendicular to the bulkhead 8 when the landing gear 43 is fullyextended. The oleopneumatic cylinder 610 may be a hydraulic piston orair filled piston. The oleopneumatic cylinder 610 has a first endconnected to the swing arm 604 between the wheel 602 and the forwardlink arm 606. In one embodiment, the oleopneumatic cylinder 610 isconnected at approximately the center of the swing arm 604. Theoleopneumatic cylinder 610 passes through the plate 608 allowing thesecond end of the oleopneumatic cylinder 610 to rotatively connect tothe bulkhead 8 such that the oleopneumatic cylinder 610 rotates towardsthe bulkhead 8 as the landing gear 43 is retracted. The forward link arm606 is rotatively connected to the oleopneumatic cylinder 610 at apositioned just below the plate 608. The actuation device 612 isrotatively coupled to the bulkhead 8 by a hinge and to the plate 608 bya hinge. The actuation device 612 includes a base portion 614. Theactuation device 602 may be a hydraulic actuator, a linear actuator orany other device capable of retracting and extending the landing gear43.

FIG. 6C depicts the landing gear 43 as the landing gear 43 is retractedinto the fuselage. The actuation device 612 is activated such that theextension arm 614 retracts into the actuation device 612 pulling theplate 608 towards the bulkhead 8. As the plate rotates towards thebulkhead 8, the forward link arm 606 rotates towards the plate 608 andthe swing arm 604 rotates towards the forward link arm 606 pulling thewheel 602 upward. FIG. 6D depicts the landing gear 43 retracting intothe fuselage. As the actuation device 612 continues to pull theextension arm 614 into the base 600, the plate 608 is pulled furthertowards the bulkhead 8 causing the oleopneumatic cylinder 610 to rotateupward and compress, and the extension arms 604 and 606 rotates towardsthe plate 608 pulling the wheel 602 upwards into the fuselage. FIG. 6Edepicts the landing gear 43 fully retracted into the fuselage. Thelanding gear is extended by extending the extension arm 614 out of theactuation device 612 such that the plate 608 rotates away from thebulkhead 8.

FIG. 7A depicts the rear landing gear 700 connected to truss element 20.The rear landing gear 700 includes two frames 702 that are eachsubstantially A-shaped. Each frame 716 is rotatively affixed to a sideof the truss element 20 by a pin. Each frame 716 is also rotativelyconnected to a trailing link 704 by a pivot joint 706. The pivot joint706 is substantially ‘C’ shaped and is sized to accommodate an end ofthe trailing link 704. A pin 708 passes through both sides of the pivotjoint 706 and the trailing link 704 to secure the trailing link 704 inthe pivot joint 706. The opposite end of the trailing link 704 isconnected to a wheel 710 and one end of a cylinder 712. The other end ofeach cylinder 712 is rotatively connected to a support unit 553 on thetruss element 20 via a universal joint.

Each frame 702 includes an overcenter locking unit 714 that isconfigured to secure the frame in a fully extended position and asupport plate 716 rotatively connected to the truss element 20 by ahinge. The end of the locking unit 714 furthest from the truss element20 is rotatively coupled to the end of the support plate 716 furthestfrom the truss element 20. Each locking unit 714 is separated into twosections by a pin. The cylinder 712 may be an hydraulic piston filledwith a hydraulic fluid and air. The cylinder 712 includes a cylinderbody 718 and rod 720 extending from the cylinder body 718.

FIG. 7B depicts the rear landing gear 700 in the fully extendedposition. The locking units 714 are fully extended such that the supportplate 716 is substantially perpendicular to the side of the trusselement 20. FIG. 7C depicts the rear landing gear 700 retracting intothe fuselage. A retraction cylinder folds the locking units 714, pullingthe support plate 716 upward. As the support plate 716 moves upward, thetwo portions of the locking unit 714 rotate about the pin, separatingthe two portions of the locking unit 714 such that the two portions ofthe locking unit 714 move towards each other. The movement of thecylinder 712 causes the trailing link 704 to rotate towards the trusselement 20, bringing the wheels 710 towards the fuselage. FIG. 7Ddepicts the rear landing gear 700 further retracting into the fuselage.As the support plate 716 continues to move towards the truss element 20,the cylinder pulls the wheels 710 into the fuselage.

FIG. 7E depicts the rear landing gear 700 fully refracted into thefuselage. The rod 720 is fully extended out of cylinder 712, and thesupport plate 716 and the central axis of the wheel 710 both aresubstantially parallel to the side of the truss element 20. The twoportions of the locking unit 714 are separated by an angle with theangle being less than 90 degrees.

Propulsion of the aircraft may be provided by a fixed-pitch eight bladecomposite blade propeller mounted at the rear of the fuselage on thecenterline axis. The propeller airfoil sections and section incidenceangles are configured to provide maximum efficiency at cruise at 50,000ft. altitude and above. Propeller diameter is also optimized for thehigh altitude cruise environment and as a result essentially eliminatessupersonic blade velocities during low altitude operation. The optimumpropeller diameter is slightly smaller than maximum fuselage diameterwhich coincidentally reduces the probability of bird strike and otherforeign object damage.

The propeller is connected to two engines by a drive shaft extendingfrom the output shaft of a gear box. The engines are liquid-cooleddiesel engines driving torque converters connected to the gear box.Multi-stage turbo charging is provided to compensate for altitude and toprovide cabin pressurization. Engine heat exchangers, turbo chargers andintercooler heat exchangers are all mounted in ducts configured toprovide thermal recovery of waste heat for supplemental propulsion.Engine exhaust is likewise used in the rear of the same duct to providean injection pump function both for cooling air circulation during lowspeed operation and to provide additional thrust during flight.

The torque converters are provided to isolate the propeller, driveshaft, and gear box from periodic variations of engine torque and toprovide for necessary torque multiplication required by the propellerduring low speed operations. Traditional propeller and enginecombinations provide no vibration isolation and match engine torqueoutput to propeller demands by varying the pitch of the propeller toreduce the propeller torque demand. This results in much higherpropeller speeds during near ground operations, and consequently muchgreater noise output, and it also results in a propeller airfoil andpitch distribution that is never optimum. The use of torque converterswithout lockup clutches allows an engine shutdown to disconnect theinoperative engine from the driveshaft and propeller. In the event thatboth engines are shut down, the propeller is completely disconnectedfrom both engines. Alternators and emergency cabin pressurization remainconnected to the drive shaft and are driven by the wind millingpropeller. This is the only external mechanical drag load applied to thepropeller aside from bearing friction and freewheeling transmissionfriction.

FIG. 8 depicts a heat recovery system 800 used to increase theefficiency of the aircraft. Cooling air is introduced to the heatrecovery system 800 from ducts 802 located on the exterior of theaircraft. The ducts may be NACA submerged ducts. The air introduced viathe ducts 802 passes over a first heat exchanger 804. The first heatexchanger 804 provides cool fluid used to cool the air bled from theturbo charger used to pressurize the cabin. The air then passes over asecond heat exchanger 806 that provides cooling liquid for theintercoolers that cool the engine air intake. The air then passesthrough a third heat exchanger 808 that cools the liquid from the enginejacket.

After leaving the third heat exchanger 808, the air passes across theturbo chargers 810. The output of the turbo chargers 810 are connectedto the manifold 812 and intercoolers of the engine to provide compressedair to the engine to increase the thrust produced by the engine. Theturbine exhaust of all turbo chargers on each side are combined into asingle tubular exhaust pipe 814 which combines with a convergent part ofthe duct 816 to form an injection pump that mixes the turbine exhaustwith the heated cooling air flow and then flows through a nozzle toprovide additional thrust. In one embodiment, the thermal recoverysystem 800 generates an additional 5-6 pounds of thrust.

FIG. 9 depicts a side view of the wing spar 900 of the aircraft ofFIG. 1. The wing skin 902 and a sleeve 904 are bonded to the skin 902 atupper and lower surfaces and at corners 906 of the sleeve 904. Thesleeve 904 is a tight fit to the wing spar 900 and is pinned to the spar900 at the wing root by a pin located on the neutral axis of the spar900. A spoiler 910 and vent 912 are provided for roll control and flightpath control. The spoiler 910 and vent 912 are linked to open togetherto provide a slot lip type aileron. The wing skin 902 is bondedinternally to the sleeve 904 such that the skin 904 that slips over theoutside of the spar 900 to form a close fit to the spar 900 that is freeto slide in the span wise direction to accommodate flexure of the spar900. In one embodiment, the skin 902 is fastened to the spar 900 at thewing root only. By securing the skin 902 to the spar 900 at the wingroot only, the skin 902 is isolated from the spar 900 in order tominimize skin 902 buckling due to bending and to allow for quickreplacement of damaged skin sections 902, ease of updating of wingsystems and airfoil shapes, and quick installation and removal forinspection of the spar 900 structure and the flap and spoiler systems.

FIG. 10A depicts a flap control system 1000 included in the wing of theaircraft in FIG. 1. The flap control system 1000 includes a plurality ofcontrol stations 1001 that each includes a plurality of plates 1002,1004, 1006 and 1008 connected together by fasteners 1010 passing throughthe corners of each plate. Each plate 1002, 1004, 1006 and 1008 includesan opening 1012 that is sized to accommodate a drive shaft 1014. Eachstation 1001 is secured to the wing spar 900. The drive shaft 1014extends the length of the wing and is connected to each control station1001. The plates 1004 and 1006 have a length longer than the plates 1002and 1008. One end of the plates 1004 and 1006 includes an opening 1016that is sized and shaped to accommodate a fore flap 1018. The fore flap1018 is connected to a flap 1020 by a flap plate (not shown).

FIG. 10B depicts the flap control system 1000 with plates 1002, 1004 and1008 removed. A chain 1050 is driven by the drive shaft 1014 connectedto a sprocket 1052, and which wraps around idler gears 1054 and 1056.The drive shaft 1014 rotates both clockwise and counterclockwise todrive the chain 1050 in both forward and reverse directions to extendand retract the flap 1020. The chain 1050 is tensioned by the idlergears 1054 and 1056 and is attached to chain shoe 1058. The chain shoe1058 is positioned and slides in slot 1060 on the inner surface of plate1008 and is rotatively connected to one end of a support arm 1062 suchthat the chain shoe 1058 can rotate relative to the support arm 1062.The opposite end of the support arm 1062 connects to the foreflap 1018through a slot 1064. A second shoe 1068 is connected to the support arm1062 at approximately the center of the support arm 1062. The secondshoe 1068 is positioned and slides in slot 1070 in plate 1008. Slot 1070is substantially arc shaped and is positioned to allow optimumpositioning of the flap 1020 or foreflap 1018 with respect to the wing.A link arm 1072 is substantially ‘U’ shaped and is connected to thesecond shoe 1068 at substantially the center of the link arm 1072. Oneend of the link arm 1072 is coupled to a third shoe 1074 that ispositioned and slides in a slot 1076. Slot 1076 is substantially arcshaped and is positioned below the slot 1070. The end of the link arm1072 opposite the end connected to the third shoe 1074 is connected totilt arm 1078. The end of the tilt arm 1078 not connected to the linkarm 1072 is connected to the lower portion of the flap plate 1066 at aposition below the connection of the support arm 1068 to the flap plate1066.

FIG. 11A depicts the flap control system 1000 in the retracted or zerodegree position. The chain shoe 1058 is positioned adjacent to the idlergear 1056 in the slot 1060, the third shoe 1074 is positioned near thebottom edge of the plate 1008 in the slot 1076 and the tilt arm 1078 isin its full refracted position. FIG. 11B depicts the flap control system1000 extending the flap 1020 downward. The sprocket 1052 drives thechain 1060 moving the sprocket 1052 towards the flap 1020. As thesprocket 1058 moves, the support arm 1062 pushes the foreflap 1018 andthe flap 1020 outwards. As the support arm 1062 moves, the link arm 1072moves in the slot 1076 pulling the tilt arm 1078 inwards causing theflap plate 1066 to rotate in a clockwise manner.

FIG. 11C depicts the flap control system 1000 as it extends furtheroutwards. As the chain 1050 continues to move the chain shoe 1058 thesupport arm 1062 pushes and rotates the foreflap 1018 and the link arm1072 continues to move in the slot 1076 to push the tilt arm 1078 awayfrom the plate 1008 to rotate the foreflap 1018 and flap 1020 down. FIG.11D depicts the flap control system with the foreflap 1018 and flap 1020in the full extended position. The chain shoe 1058 is positioned in theportion of the slot 1060 furthest outward. The link arm 1072 ispositioned in the slot 1076 such that a portion of the link arm 1072 issubstantially perpendicular to the tilt arm 1078. The flap 1020 ispositioned such that the training edge of the flap 1020 pointssubstantially downward.

The flap control system may be a 90% span double-slotted flap systemincluding slot lip spoilers and spoiler vents used for roll control andglide path modulation. All flap tracks are fully internal to the wingwhen the flaps are refracted, and extension is by means of drive shaft1014 extending across the full 90% of span with the drive shaft actuatorin the center of the wing. Each control station 1001 along the wingconverts rotational motion of the drive shaft 1014 to linear motion ofthe support arm 1062 and the link arm 1072 and the motion of the tiltarm 1078 by means of the sprocket 1056 and chain 1050. The tooth countof each sprocket 1056 is a fixed ratio to chord length of the wing ateach span wise station.

FIG. 12 depicts the spoiler actuation system 1200 used to actuate thespoiler 91 of FIG. 9. The spoiler 910 is actuated by means of twoslotted mount plates 1202 and 1204 plates and a cam plate 1205 toprovide positive control of extension and retraction of the spoiler 91and full lock of the spoiler 910 in the refracted position. Normally,the cam plates 1205 are linked together and move synchronously, lockingone spoiler in the locked down position while proportionately deployingthe opposite spoiler with respect to the yoke rotation. Approach pathmodulation is provided by moving the cam plates 1205 on opposite wingseither closer together or farther apart with respect to one another. Theentire flap and spoiler mechanism is mounted in the open rear half ofthe spar of the wing, which provides unrestricted access to themechanism when the wing glove is removed.

FIG. 13A depicts a trim actuator 1300 that is mechanically coupled tothe elevator control system. A similar actuator is used on the dorsalfin control system. The actuator includes a base housing 1302 and anextension rod 1304 that slides into and out of the base housing 1302.The end of the extension rod 1304 opposite the base housing 1302 and theend of the base housing 1302 opposite the extension rod 1304 eachincludes a securing unit 1306 and 1308 affixed to the end thereon. Thesecuring units 1306 and 1308 may be eyelets.

FIG. 13B depicts an interior view of the actuator 1300 in a centered,compressed and extended position. The base housing 1302 contains twosprings 1310 and 1312 and a stop 1314 fastened to the cylinder bore. Theextension rod piston 1304 engages two washers 1316 and 1318 that lie oneither side of the stop 1314 and against which the springs 1310 and 1312rest. When the extension piston 1304 is moved in either direction fromits neutral position aligned with the stop 1314, it compresses one ofthe springs 1310 and 1312 which drives the extension rod piston 1304back into the neutral position. The overall position of the actuator iscontrolled by a ball bearing jack screw that sets the trim position ofthe elevator, and a second similar system sets the position of thedorsal fin. The surfaces of the extension rod 1304 and base housing 1302are never in a stick-free condition, thus eliminating the need forgeared tabs and other complications for stabilization.

The aircraft cabin may be approximately 74 inches high and include anapproximately 78 inch width having a minimum 50 inch seat pitch. Theaircraft has a service ceiling of approximately 65,000 feet, and anormal cruise speed of between approximately 460 to approximately 510mph, with a specific fuel consumption of approximately 30 toapproximately 42 mpg depending on cruise speed and altitude. Landingstall speed is approximately 70 mph, takeoff and landing speeds areapproximately 90 mph, and runway requirements are approximately 3000 ft.

It is to be understood that both the foregoing general description andthe following detailed description are exemplary and are intended toprovide further explanation of the invention claimed. The disclosedconfiguration is the preferred embodiment and is not intended topreclude functional equivalents to the various elements.

The accompanying drawings, which are incorporated in and constitute partof this specification, are included to illustrate and provide a furtherunderstanding of the invention. Together with the description, thedrawings serve to explain the principles of the invention.

What is claimed:
 1. A supplemental thrust system for an airplanecomprising: a submerged inlet duct; a diffuser section in communicationwith the inlet duct; and a duct outlet in communication with thediffuser section; an outlet nozzle in communication with the ductoutlet; a turbocharger unit comprising one or more turbocharges incommunication with each other to provide compressed air to the engineand the cabin; a heat exchanger unit comprising a plurality of heatexchangers in the diffuser section to provide cooling liquid to coolpressurized air in the cabin, engine intake air and the engine jacketwherein each turbocharger is coupled to the intake manifold of an engineto increase the power of the engine by introducing compressed air intothe manifold.
 2. The supplemental thrust system of claim 1, whereinturbine exhaust of each of the turbochargers converges into a singleduct that is connected to the duct outlet.
 3. The supplemental thrustsystem of claim 1, including an expansion duct between the duct outletand the outlet nozzle.
 4. The supplemental thrust system of claim 1,wherein the inlet duct is a NACA duct on the exterior of an aircraft.